This invention relates to a thermal barrier coating system having enhanced resistance to spallation, and, more particularly, to such a system wherein adhesion is enhanced and crack propagation is reduced by physical modification of at least one surface underlying the ceramic thermal barrier coating.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of hot gas turns the turbine, which turns the shaft and provides power to the compressor. The hot exhaust gases then flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the exhaust gas temperature. However, the maximum temperature of the exhaust gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of superalloys, and can operate at temperatures of up to about 1900.degree.-2100.degree. F. As used herein, the term superalloy includes high-temperature-resistant alloys based on nickel, cobalt, iron or combinations thereof.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes, and other components of the engine that operate at high temperatures. The composition and processing of the materials themselves have been improved. Physical cooling techniques are used. In one widely used approach, internal cooling channels are provided within the components, and cool air is forced through the channels during engine operation.
To provide, a further increase in the operating temperature limit, a thermal barrier coating system is applied to the turbine blade or turbine vane, which acts as a substrate. The thermal barrier coating system includes a ceramic thermal barrier coating that insulates the component from the hot exhaust gas, permitting the exhaust gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component. Ceramic thermal barrier coatings usually do not adhere well directly to the superalloys used in the substrates, and therefore an additional layer called a bond coat is typically placed between the substrate and the thermal barrier coating. The bond coat improves the adhesion, and, depending upon its composition and processing, may also improve oxidation and corrosion resistance of the substrate.
The thermal barrier coating system must remain in place on the protected component to be useful. When the component is repeatedly heated and cooled, as occurs in the operating cycles of the gas turbine engine, thermally induced stresses and strains are produced and accumulate within the thermal barrier coating system due to the different thermal expansion coefficients of the ceramic thermal barrier coating and the metallic substrate to which it is applied. The bond coat helps to alleviate the buildup of stresses and strains, but they are present. The bond coat also improves the adhesion of the thermal barrier coating by improving the oxidation resistance of the substrate.
The most common mechanism of failure of the thermal barrier coating system is the spallation of the coating in local regions of the protected component. A crack is produced in the thermal barrier coating due to the accumulation of stresses and strains. The crack eventually propagates until a portion of the coating system flakes or chips away, this process being termed "spallation". Such spallation failure usually occurs in patches. With the thermal barrier coating system locally removed, the underlying component is exposed to the hot exhaust gas temperatures, above which the unprotected component can not operate, and failure of the component quickly follows.
A number of techniques have been developed to reduce the tendency toward spallation failure of the thermal barrier coating. These techniques include optimization of compositions of the various layers, optimization of processing, adding new layers to the bond coat, and changes in design of the underlying components. The various approaches have been successful to varying degrees, but also involve drawbacks such as increased weight, constraints on design, and manufacturing complexity. Although progress has been made, the problem of spallation failure of thermal barrier coating systems remains.
There is therefore a need for an improved approach to improving the resistance to spallation failure of components protected by thermal barrier coating systems. The approach should be operable to extend the life of the protected component, and should be compatible with commercial production of engine components. The present invention fulfills this need, and further provides related advantages.